The present invention relates to spacecraft and solid rocket vehicles and, more particularly, to control of disturbance torques during solid rocket burn for a spinning vehicle.
It is conventional to use multiple-stage rockets for putting artificial satellites into earth orbit. Generally, two stages are used for the initial boost phase to achieve a low orbit outside the Earth's atmosphere. A third stage produces a highly elliptical orbit known as the transfer orbit, and a fourth stage is used when a high altitude circular orbit is required. A solid rocket motor in a spinning vehicle configuration is common for the third and fourth stages, at spin rates of 20 to 60 rpm, typically.
During firing of a spinning solid rocket motor, the spacecraft is often subjected to disturbance torques that cause a coning motion known as nutation. As long as the nutation is small, it is not objectionable because it can be controlled or damped, and there is no danger of the satellite tumbling. However, if the nutation increases to the range between 8 to 15 degrees, there is cause for alarm.
This coning motion is thought to be caused by the motion of the combustion products within the solid rocket motor during firing. One method of controlling this coning is by putting constraints on the manufacturer of the solid rocket motor and on the manufacturer of the spacecraft. The spacecraft mass properties and the solid rocket motor lateral force characteristics can be limited and specified in great detail to prevent or eliminate disturbance torques that cause coning motion.
Another method of compensating for the coning motion is to provide dedicated control and propulsion equipment such as steering nozzles integral with the solid rocket motor that are automatically aimed in a direction that compensates for lateral disturbance torques. Another method is to provide a self-contained external bolt-on package that adds dedicated control and propulsion thrusters to the solid rocket motor for compensating lateral disturbance torques automatically.
It is an objective of the present invention to provide a solution to the problem of nutation during firing of a spinning solid rocket vehicle that also reduces or avoids limitations or constraints on spacecraft mass properties and solid rocket motor lateral force characteristics. Another objective of the present invention is the provision of a solution to the problem of nutation during firing of a spinning solid rocket vehicle that does not require the use of dedicated control and propulsion equipment. Yet another objective of the invention is to provide a solution to the problem of nutation during firing of a spinning solid rocket vehicle that is more efficient in weight and cost than alternative solutions such as those referred to above.